A gas turbine engine such as that configured for powering an aircraft in flight conventionally includes in serial flow communication a fan, a compressor, a combustor, a high pressure turbine (HPT), and a low pressure or power turbine (LPT). Ambient air enters the fan wherein it is initially pressurized, and in turn a portion thereof flows to the compressor wherein it is further pressurized and discharged to the combustor wherein it is mixed with fuel and ignited for generating hot combustion gases which flow downstream to the HPT. The HPT includes one or more stages of turbine blades specifically configured for extracting energy from the combustion gases for powering the compressor through a shaft connected therebetween. The combustion gases lose pressure in the HPT and then flow to the LPT which includes additional turbine blades also configured for extracting additional energy from the lower pressure combustion gases for powering the fan connected thereto by another shaft.
The fan and compressor include respective rotor blades which are configured for pressurizing the relatively cool air which is in contrast to the turbine blades of the HPT and the LPT which are configured for extracting energy from the hot combustion gases with a resulting reduction in pressure thereof. The energy extracted from the combustion gases is in turn imparted to the air being pressurized in the fan and compressor.
A fan blade as used herein is merely a type of generic compressor blade since both blades are effective for imparting energy into the air for increasing its pressure to different levels. The fan blade is relatively large for moving larger amounts of airflow at reduced pressure for providing a substantial portion of propulsion thrust from the engine. The fan blades are typically configured in one or two stages for use in conventional high bypass, turbofan, commercial aircraft engines or lower bypass military engines.
The rotor blades found in a typical axial compressor are configured in a substantial number of axial stages with each succeeding stage having smaller and smaller rotor blades for incrementally increasing pressure of the airflow channeled therethrough.
Accordingly, compressor and turbine blades are fundamentally different, and each is designed for maximizing aerodynamic efficiency at predetermined design speeds of operation such as the maximum speed associated with takeoff operation of an aircraft or the lower speed associated with cruise operation of the aircraft. Compressor blades, and fan blades as used herein, must also be designed to provide an adequate stall margin at maximum compression ratios for maximizing propulsion thrust with minimum weight and fuel consumption. Since the compressor blades pressurize the air used in the combustion process for generating the combustion gases, they affect the temperature of the combustion gases discharged to the HPT which must remain within acceptable levels for ensuring a useful life of the HPT rotor blades.
Furthermore, fan and compressor blades may be operated at various speeds with the airflow thereover varying in speed from subsonic, transonic, and supersonic which also must be accommodated in blade design. In a fan blade, for example, the blades generate noise during operation which must be maintained within acceptable limits.
Accordingly, improvement in fan or compressor blade performance is desirable for increasing efficiency and stall margin at increased stage compression ratios. In turn, thrust may be increased with reductions in weight and fuel consumption, and fewer stages may be used in a typical compressor. Reduced aeromechanical excitation and improved noise characteristics may also be obtained by improving fan blade design. And, lower turbine temperatures may also be obtained from improved fan and compressor blade designs for increasing hot section life.